1. Field of the Invention
The present invention relates to an airfoil which is suitably used for a blade cascade of an axial-flow compressor for transonic velocity of an aircraft engine. More particularly, to an airfoil that is capable of a drastic reduction of a pressure loss in a low Reynolds number region not more than a critical Reynolds number that corresponds to a starting point below which the total pressure losses increase considerably.
2. Description of Background Art
Currently, as an airfoil is known that is widely used in a blade cascade (rotor blade, stator vane, outlet guide vane) for an axial-flow compressor for small to large-size state-of-the-art aircraft engines, Controlled Diffusion Airfoil (CDA). In this CDA, a maximum flow velocity on an extrados of a blade in a transonic regime is generated over a portion of the suction surface from 10 to 30% of a chord. A concept of its design is to provide a flow velocity distribution wherein the flow velocity is reduced from supersonic to subsonic without a shock wave so that shock losses are eliminated and the boundary layer is not separated due to shock-boundary layer interaction.
Japanese Patent Application Laid-open No. 2002-317797 discloses an airfoil in which a surface having a surface roughness that is relatively larger on a front half part of a portion from a leading edge to an extrados than a rear half part is formed on the airfoil so as to suppress the generation of laminar flow separation bubbles and to suppress the development of a turbulent boundary layer in a low Reynolds number region as well as to prevent a decrease in a surge allowance, thereby improving efficiency of the compressor.
Also, Japanese Patent Application Laid-open No. 2004-293335 discloses an airfoil in which a supersonic portion having a substantially constant flow velocity is formed in a region downstream of a first maximum flow velocity value on an extrados of an airfoil for a compressor and within 15% on a chord so that a large first shock wave is generated at a position where the flow velocity becomes the first maximum value, thereby weakening a second shock wave generated at a position where the flow velocity becomes substantially a constant supersonic velocity. Thus, a boundary layer separation due to the second shock wave is suppressed to reduce a pressure loss.
It is very important for an aircraft engine to reduce the weight. The weight of LP turbine accounts for roughly one third of the total engine weight, because it consists of several stages. An idea of reducing the number of turbine components is to include a high turning compressor stator as an outlet guide vane (OGV) just behind an extremely high loaded turbine rotor. However, the operating Reynolds number varies greatly between take-off and cruise condition. As a result, airfoils of conventional medium and high Reynolds number CDA design do have problems at cruise conditions at a low Reynolds number region less than a critical Reynolds number. Indeed, the OGV losses could dramatically increase below a certain Reynolds number, so that a sufficient performance of the aero engine cannot be achieved.
Total pressure losses of conventional aero engine compressor bladings also increase tremendously at very high altitude cruise (i.e. above 40000-45000 ft) at which the blade chord Reynolds number is very low because of the low air density.